Hypersonic missiles have a future Naval need to reduce the time to impact on time critical targets. Supersonic combustion is a very difficult subject that has been attacked often in the past with limited success. Hypersonic missiles have utilized both ramjet and scramjet technologies and designs to reach both high speeds and long-range capabilities.
FIG. 4A illustrates an example of a ramjet engine design which operates by subsonic combustion of fuel in a stream of air compressed by the forward speed of the aircraft itself, as opposed to a normal jet engine, in which the compressor section (the fan blades) compresses the air. Ramjets operate from about Mach 2 to Mach 5. Scramjet is an acronym for Supersonic Combustion Ramjet. The scramjet differs from the ramjet in that combustion takes place at supersonic air velocities through the engine. It is mechanically simple having a burner (2), but vastly more complex aerodynamically than a jet engine. Hydrogen is normally the preferred fuel used.
A ramjet has no moving parts and achieves compression of intake air by the forward speed of the air vehicle. Air entering the intake of a supersonic aircraft is slowed by aerodynamic diffusion created by the inlet and diffuser (1) to velocities comparable to those in a turbojet augmentor. The expansion of hot gases after fuel injection and combustion accelerates the exhaust air to a velocity higher than that at the inlet and creates positive push. Solid fuel ramjet engines, whether brought to operational speed by a booster engine or air dropped from a vehicle, depend upon the introduction of air into the engine due to its forward motion. Thus the term ramjet is used. As the ram air passes through a solid fuel grain within a combustor, fuel rich gases generated by the solid fuel react with oxygen in the air inside the solid fuel bore and in the further downstream located mixing chamber of the combustor and pass out of the engine via a nozzle (3) producing thrust.
FIG. 4B illustrates an example of a scramjet engine design that (supersonic-combustion ramjet) is a ramjet engine in which the airflow through the whole engine remains supersonic. The scramjet has an inlet (1), burner (2), and nozzle (3). Scramjet technology is challenging because only limited testing can be performed in ground facilities. A scramjet works by taking in air at speeds greater than Mach 1, and using it to ignite pollution-free hydrogen, which in turn creates super-propulsion.
The speed of sound is generally about 1,300 kilometers per hour, and supersonic flight is deemed to be anything between that and Mach 4, or four times the speed of sound. Hypersonic speeds lie above that. The Concorde flies at Mach 2.2. The fastest current existing air-breathing jet, known as the SR-71 Blackbird, flies at Mach 3.6. For example, at Mach 10—or 10 times the speed of sound—the 12-foot-long, 5-foot-wide aircraft will be travelling at about two miles per second (approximately 7,200 miles per hour at sea level). Speeds over Mach 5 are defined as “hypersonic.” (The Aviation History On-line Museum & GE Aircraft Engines).
Due to a wide range of flight conditions encountered by these engines during operation, the air mass flow varies considerably while the missile is changing speed and altitude. Without some means of controlling the burn rate of the solid fuel in response to changes in air mass flow excessively rich combustion chamber conditions will exist, which is very wasteful of fuel and reduces the range of the vehicle. Additionally, engine variables, such as changes in the solid fuel grain area, thrust, and combustor temperatures and pressures, as well as missile flight parameters, such as Mach number and angle of attack necessitate changes in fuel burn rate to maintain the variable within acceptable limits.
Combustion instability has been a problem of major concern. Unstable, periodic fluctuation of combustion chamber pressure that has been encountered in ramburners arises from several causes associated with combustion mechanism, aerodynamic conditions, real or apparent shifts in fuel-to-air ratio or heat release, and acoustic resonance. The periodic shedding of vortices produced in highly sheared gas flows has been recognized as a source of substantial acoustic energy for many years. For example, experimental studies have demonstrated that vortex shedding from gas flow restrictors disposed in large, segmented, solid propellant rocket motors couples with the combustion chamber acoustics to generate substantial acoustic pressures. The maximum acoustic energies are produced when the vortex shedding frequency matches one of the acoustic resonances of the combustor. It has been demonstrated that locating the restrictors near a velocity antinode generated the maximum acoustic pressures in a solid propellant rocket motor, with a highly sheared flow occurring at the grain transition boundary in boost/sustain type tactical solid propellant rocket motors.
An apparatus and method for controlling pressure oscillations caused by vortex shedding is disclosed is in U.S. Pat. No. 4,760,695 issued to Brown, et al. on Aug. 2, 1988. The '695 patent discloses an apparatus and method for controlling pressure oscillations caused by vortex shedding. Vortex shedding can lead to excessive thrust oscillations and motor vibrations, having a detrimental effect on performance. This is achieved by restricting the grain transition boundary or combustor inlet at the sudden expansion geometry, such that the gas flow separates upstream and produces a vena contracta downstream of the restriction, which combine to preclude the formation of acoustic pressure instabilities in the flowing gas stream. Such an inlet restriction also inhibits the feedback of acoustic pressure to the point of upstream gas flow separation, thereby preventing the formation of organized oscillations. The '695 patent does not present a method or apparatus, which attempts to permit a significant portion of the required enthalpy proportioned to an expansion side of the nozzle via supersonic combustion without the use of expensive film cooled nozzles. Furthermore, the '695 patent does not utilize an oxygen injection means for maintaining flame stability.
With long-duration hypersonic flight come material problems. The conventional approach to creating these high Mach high enthalpy flows is to expand very high temperature combustion through a nozzle to the desired pressure, temperature, and Mach. However, the high total temperature required puts extreme erosion on the throat of the nozzle. As a result, the conventional high temperature subsonic combustion and nozzle expansion approach requires the use of complex and expensive film cooled nozzles (estimated to be at the cost of $2 million) to survive the high enthalpy flow conditions for the relatively long test times required by the use of such device.
Therefore, there remains a need to develop a supersonic combustion heater that enhances kinetics, produces an increased high enthalpy flow source, enhances flame stability, improves mixing between fuel and air, and shortens chemical ignition delay, without the use of expensive film cooled nozzles.